The invention relates to the cooling of components of a gas turbine engine; and in particular, heat exchangers utilizing bypass air to cool such components.
In a gas turbine engine such as those typically used to power aircraft, a compressor discharges high pressure compressed air into a combustion chamber where it is mixed with fuel and the mixture burned. The resulting combustion gases drive a turbine which provides engine power and thrust. In order to increase the efficiency and power of the turbine, it is desirable that the combustion gases have the highest possible temperature. On the other hand, it is desirable that the turbine components driven by the exhaust gases be maintained at a lower temperature than the combustion gases to prevent degradation of the components.
Gas turbine engines have been provided with mechanisms for cooling the turbine components, thereby protecting the components from the extreme heat of the exhaust gases. Techniques for providing cooling air through and around turbine blades are well known. For example, U.S. Pat. No. 2,479,573 issued to A. Howard illustrates a cooling mechanism which provides a film of cool air over the turbine blades.
Bypass air supplied by a fan at the air inlet end of the engine is substantially cooler than the air being worked within the engine. It is known to use bypass air to cool compressed air in an air-to-air heat exchanger located in the bypass air duct. The heat exchanger receives compressed air from the core, and cooled air is returned to the core as secondary cooling air and routed to turbine components for more effective cooling.
Typically, a standard serpentine flow heat exchanger is used in the bypass duct for cooling secondary air, but such a device has several disadvantages in this environment. For example, such heat exchangers typically are relatively bulky and tend to negate any other weight reduction attempts within the engine. Since such heat exchangers are manifested by one or more large box-like mechanisms equally spaced around the circumference of the bypass air duct, they disrupt the uniform flow of bypass air through the bypass duct which may lead to downstream hot spots. Further, the serpentine air path of such heat exchangers presents a substantial resistance to the cooling air flow and results in a significant pressure drop as the air travels through the heat exchanger.
Such systems also require manifolds for drawing and returning the air to prevent localized pressure effects from inhibiting combustor or turbine performance. In order to absorb localized thermal-growth differences between one end of the heat exchanger located at the high temperature turbine inlet and the other end of the heat exchanger located at the cooler compressor discharge, special seal mechanisms are required which allow relative motion while simultaneously being subjected to high pressure. Accordingly, there is a need for a heat exchanger for use with a turbofan engine which minimizes any disruption of bypass air flow and does not require expensive seals to compensate for dimensional variations brought about by thermal gradients.